The present invention relates to components having internal passages through which cooling air flows and is then discharged through surface openings to provide an air film cooling effect. More particularly, this invention is directed to a gas turbine engine airfoil equipped with a diffuser cooling hole whose configuration increases the effectiveness of the cooling film.
The operating environment within a gas turbine engine is both thermally and chemically hostile. Significant advances in high temperature alloys have been achieved through the formulation of iron, nickel and cobalt-base superalloys, though components formed from such alloys often cannot withstand long service exposures if located in certain sections of a gas turbine engine, such as the turbine, combustor or augmentor. A common solution is to protect the surfaces of such components with an environmental coating system, such as an aluminide coating or a thermal barrier coating (TBC) system. The latter typically includes an environmentally-resistant bond coat and a thermal barrier coating of ceramic deposited on the bond coat. Bond coats are typically formed from an oxidation-resistant alloy such as MCrAlY where M is iron, cobalt and/or nickel, or from a diffusion aluminide or platinum aluminide that forms an oxidation-resistant intermetallic. During high temperature excursions, these bond coats form an oxide layer or scale that bonds the ceramic layer to the bond coat. Zirconia (ZrO2) that is partially or fully stabilized by yttria (Y2O3), magnesia (MgO) or other oxides has been widely employed as the material for the ceramic layer. The ceramic layer is typically deposited by air plasma spraying (APS), low pressure plasma spraying (LPPS), or a physical vapor deposition (PVD) technique, such as electron beam physical vapor deposition (EBPVD) which yields a strain-tolerant columnar grain structure.
While thermal barrier coating systems provide significant thermal protection to the underlying component substrate, internal cooling of components such as turbine blades (buckets) and nozzles (vanes) is generally necessary, and may be employed in combination with or in lieu of a thermal barrier coating. For this purpose, airfoils of turbine blades and nozzles of a gas turbine engine often require a complex cooling scheme in which bleed air is forced through passages within the airfoil and then discharged from the airfoil through carefully configured cooling holes. As an example, FIG. 1 shows an airfoil 110 having a cooling cavity 112 and a diffuser cooling hole 114. The cooling hole 114 has a divergent opening 116 at the exterior surface 117 of the airfoil 110 to promote the distribution of a cooling film over the downstream airfoil contour and therefore increase the effectiveness of the cooling film. The performance of a turbine airfoil is directly related to the ability to provide uniform cooling of its surfaces with a limited amount of cooling air. In particular, the size and shape of each opening determine the amount of air flow exiting the opening and the distribution of the air flow across the downstream surface of the airfoil, and also affect the overall flow distribution within the cooling circuit containing the opening. Other factors, such as backflow margin, are also affected by variations in opening size. Consequently, cooling holes, their openings and the processes by which they are formed and configured are critical.
For airfoils without a thermal barrier coating, cooling holes are typically formed by such conventional drilling techniques as electrical-discharge machining (EDM) and laser machining. An example of a diffuser cooling hole 122 formed by EDM and laser machining methods is depicted in FIG. 2. The noncircular diffuser opening 124 is generated by removing extra material at the airfoil surface along one side of the opening 122, such that the central axis 118 of the hole 122 diverges from a linear centerline 120 only in the immediate vicinity of the opening 124, but otherwise coincides with the centerline 120 throughout the remaining length of the hole 122. Notably, the lower wall of the hole (i.e., the wall farthest from the surface and on the side of the hole 122 enlarged by the opening 124) is arcuate near the surface as a result of the manner in which the opening 124 was enlarged and shaped, while the upper wall of the hole 122 (i.e., the wall nearest the surface and diametrically-opposite the side of the hole 122 enlarged by the opening 124) is substantially unaffected by the opening 124, and is therefore essentially straight. Other than in the immediate vicinity of the opening 124, the hole 122 has a roughly circular cross-section.
While EDM and laser machining techniques can be employed to produce the noncircular shape required for a diffuser opening 124, these methods are limited in their ability to tailor the shape of the cooling hole much below the airfoil surface. Another shortcoming of EDM is that cooling holes cannot be formed by this method in an airfoil having a ceramic TBC since the ceramic is electrically nonconducting. Laser machining techniques are also unacceptable for forming cooling holes in an airfoil protected by a TBC, because laser machining has a tendency to spall the brittle ceramic TBC by cracking the interface between the airfoil substrate and the ceramic. Accordingly, cooling holes have typically been formed by EDM and laser machining prior to applying the TBC system, limiting the thickness of the TBC which can be applied or necessitating a final operation to remove ceramic from the cooling holes in order to reestablish the desired size and shape of the openings.
From the above, it can be seen that the geometric configuration of a cooling hole for an air-cooled airfoil is limited by the techniques available to produce the cooling holes, particularly if the airfoil is protected by a TBC. While cooling holes formed by EDM and laser machining provide a satisfactory cooling effect, it would be desirable if the size and shape of a cooling hole could be tailored to enhance the cooling film distribution on the external surfaces of the airfoil, and thereby increase the effectiveness of the cooling film.
The present invention is embodied in a cooling hole configuration for an air-cooled component, such as a gas turbine engine airfoil. The cooling hole is configured to have cross-sectional variations and a noncircular-shaped diffuser-type opening, as well as a central axis at an acute angle to the exterior surface of the component. The cooling hole generally has a first region immediately adjacent the exterior surface, a second region beneath the first region and interior to the component, and a third region immediately adjacent to a cooling cavity within the component. A recess is present in the wall of the second region nearest the exterior surface. As a result, the recessed wall is curved instead of straight, causing the central axis of the cooling hole to have an arcuate shape in which the central axis is disposed at a lesser angle to the surface in the first region than in the second region. The recess also causes the second region of the cooling hole to have an oblong or oval cross-sectional shape as compared to the first and third regions. The first region has a larger cross-sectional area than the second region, while the third region has a circular cross-section with a smaller-cross-sectional area than the second region. Smooth transitions preferably exist between the first, second and third regions.